Missiles and rockets burn propellants within combustion chambers to generate pressurized gases. The pressurized gases are directed through a nozzle to provide thrust and accordingly propel the body of the missile or rocket. Solid rocket propellants are formed with a solid oxidizer, for instance ammonium perchlorate, plus fuels, additives and a binder ingredient. Ignition systems that elevate the temperature of the solid rocket propellant to the point of combustion are used to ignite the solid rocket fuel. After ignition of a solid rocket motor the reaction cannot be interrupted until the fuel is completely consumed. Additionally, the solid rocket propellant burns according to the shape of the propellant grain and its operating pressure which is dictated by the nozzle throat size. That is to say, burning of the fuel (the burn rate) proceeds according to a set of predefined parameters and these predefined parameters cannot be changed during launch and flight with the exception of use of a mechanical apparatus (such as a nozzle pintle). Stated another way, once the solid rocket propellant is initiated burning cannot be interrupted and there are limited means for providing launch based or in-flight control of the burn rate of the fuel.
One example of propellant configured for controllable burning in a low pressure environment (e.g., less than 200 psi) is shown in US Published Application 2011/0067789. An electrically controlled propellant is provided that is capable of sustained controllable combustion at ambient pressure. As discussed in the application at pressures of greater than 200 psi the propellant is self-sustaining. In other words, at these pressures the propellant continues to burn even with the interruption of electrical input to the propellant. In at least this regard, burning of the discussed electrically controlled propellant, like the solid rocket fuel discussed above, is not finished until the propellant is fully consumed.